XFOIL Version 6.94 Calculated polar for: LOCKHEED C-5A BL1256 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.1958 0.00708 0.00223 -0.0445 0.7606 0.7013 0.500 0.2496 0.00706 0.00232 -0.0434 0.7379 0.7599 1.000 0.3033 0.00709 0.00239 -0.0422 0.7102 0.7991 1.500 0.3579 0.00715 0.00243 -0.0413 0.6833 0.8250 2.000 0.4105 0.00725 0.00250 -0.0400 0.6490 0.8523 2.500 0.4613 0.00741 0.00259 -0.0383 0.6016 0.8777 3.000 0.5087 0.00772 0.00272 -0.0359 0.5360 0.9044 3.500 0.5521 0.00845 0.00295 -0.0332 0.4022 0.9281 4.000 0.5887 0.01017 0.00365 -0.0300 0.1754 0.9555 4.500 0.6485 0.01112 0.00425 -0.0313 0.1053 0.9860 5.000 0.7051 0.01213 0.00502 -0.0322 0.0777 1.0000 5.500 0.7581 0.01293 0.00576 -0.0321 0.0696 1.0000 6.000 0.8100 0.01375 0.00653 -0.0317 0.0640 1.0000 6.500 0.8620 0.01448 0.00727 -0.0312 0.0598 1.0000 7.000 0.9114 0.01544 0.00822 -0.0304 0.0559 1.0000 7.500 0.9592 0.01652 0.00927 -0.0294 0.0527 1.0000 8.500 1.0502 0.01923 0.01199 -0.0267 0.0467 1.0000 9.000 1.0965 0.02035 0.01327 -0.0254 0.0440 1.0000 9.500 1.1391 0.02216 0.01508 -0.0239 0.0413 1.0000 10.000 1.1819 0.02355 0.01664 -0.0222 0.0386 1.0000 10.500 1.2224 0.02566 0.01880 -0.0205 0.0365 1.0000 11.000 1.2598 0.02750 0.02091 -0.0183 0.0346 1.0000 11.500 1.2955 0.02935 0.02283 -0.0161 0.0332 1.0000 12.000 1.3219 0.03193 0.02570 -0.0128 0.0318 1.0000 12.500 1.3401 0.03408 0.02813 -0.0087 0.0304 1.0000 13.000 1.3580 0.03614 0.03033 -0.0054 0.0293 1.0000 13.500 1.3748 0.03932 0.03360 -0.0027 0.0285 1.0000 14.000 1.3709 0.04407 0.03883 0.0003 0.0280 1.0000 14.500 1.3584 0.05030 0.04553 0.0019 0.0275 1.0000 15.000 1.3367 0.05818 0.05384 0.0017 0.0270 1.0000 15.500 1.3031 0.06847 0.06457 -0.0010 0.0268 1.0000 16.000 1.2473 0.08383 0.08041 -0.0083 0.0267 1.0000